RTE is a comprehensive Rocket Thermal Evaluation computer code developed for NASA Lewis Research Center, presently Glenn Research Center. The early version of the code was distributed through NASA's COSMIC Library. The COSMIC version of RTE is marketed by AP Poineer Inc. (http://www.appioneer.com/products/11283.html). The program has been modified substantially since its early version release. This page provides a brief description of the code. For a brief slide presntation of this code with some sample results click on this text. For manual of the latest version of RTE (RTE2002) is available at the following site:
                                                http://www.tara-technologies.com/rte/rte.pdf.
If you are interested in obtaining a copy of the code please send an e-mail to: mnaraghi@tara-technologies.com.

SUMMARY

RTE (Rocket Thermal Evaluation) is  a computer code for three-dimensional thermal analysis of regeneratively cooled rocket thrust chambers and nozzles. A unique feature of this code is conjugating all thermal/fluids processes in the propulsion system in order to obtain matched results for the thermal field. These thermal/fluids processes include: convection and radiation heat transfer from hot combustion gases to the liner of the engine; conduction heat transfer with walls; and convection to the coolant. RTE uses an iterative marching scheme to match the heat flux and temperature fields of these thermal processes. The program uses GASP (GAS Properties), WASP (Water and Steam Properties) and a module for properties of RP1 to evaluate coolant flow properties. Hence, it is capable of handling all commonly used coolants in propulsion systems (e.g., H₂, O₂, H₂O, CH₄ and RP1). CET (Chemical Equilibrium with Transport Properties) code is used for evaluation of hot gas properties. The inputs to RTE consist of the composition of fuel/oxidant mixtures and flow rates, chamber pressure, coolant entrance temperature and pressure, dimensions of the engine, materials and number of nodes in different parts of the engine. It allows temperature variations in axial, radial and circumferential directions and by implementing an iterative scheme, it provides a listing of nodal temperatures, rates of heat transfer, and hot-gas and coolant thermal and transport properties. The O/F (oxidant/fuel) ratio can be varied along the thrust chamber. This feature allows the user to incorporate a non-equilibrium model or an energy release model for the hot-gas-side. The mixture ratio at each station can be calculated using ROCCID. Thermal radiation from hot gases within the chamber is also included in the analysis. The exchange factors for radiation calculations are evaluated using an external module (RTE_RAD, Rocket Thermal Evaluation Discrete Exchange Factor), which can be input to the main rocket thermal evaluation code.

This code can be used for both regeneratively and radiatively cooled engines. For regeneratively cooled engines, the code can be used for one pass as well as pass-and-half cooling cycles. Additionally, the blocked channel option allows a user to assess the thermal performance of a regeneratively cooled engine when a cooling channel is blocked. The user has the option of bypassing the hot-gas-side calculations and directly inputting gas side fluxes. This feature can be used to link RTE to a boundary layer program for the hot-gas-side heat flux calculation. The procedure for linking RTE to a hot-gas side program, TDK (Two Dimensional Kinetics Nozzle Performance Computer Program) is described in this manual.

RTE is written in Fortran and has been successfully compiled on a number of UNIX systems and Microsoft Windows. Shell programs have been developed for UNIX and WINDOWS operation systems to link RTE and TDK. To ease inputting the large data sets needed to run the program a Graphic User Interface (preprocessor) based on Excel is provided. A user can fill in engine specifications in designated Excel cells and choose the right engine information from combo boxes. Then by clicking on a command button, data from the Excel interface would be transferred into RTE’s input file.  For a trial version of RTE's GUI click on this text if you are using internet explorer, or right click on this text and save the file if you are using Netscape. Then go to the Appendix D of RTE2002 manual for instructions on using RTE's GUI. Also, RTE and its radiation module can be run from Excel. RTE provides a number of output files, each provide useful information regarding the engine’s thermal performance. The Graphic postprocessor of RTE is based on Techplot software. It produces a number of output files that can be processed by Tecplot for temperature isotherms and graphic results.

          The Rocket Thermal evaluation code is based on the geometry of a typical regeneratively-cooled engine similar to that
          shown in Figure 1.

                       Figure 1: Configuration of a typical regeneratively cooled rocket thrust chamber and nozzle

 The wall can consist of three layers: a coating, the channel, and the closeout. These three layers can be different     materials or the same material. The number of cooling channels in the wall are also specified by the user. For the numerical procedure, the rocket thrust chamber and nozzle are subdivided into a number of stations along the longitudinal direction, as shown in Figure 2.
Figure 2: Configuration of a typical regeneratively cooled thrust chamber and nozzle wall

The wall can consist of three layers: a coating, the channel, and the closeout. These three layers can be different materials or the same material. For the numerical procedure, the rocket thrust chamber and nozzle are subdivided into a number of stations along the longitudinal direction, as shown in Figure 3. The thermodynamic and transport properties of the combustion gases are evaluated using the chemical equilibrium composition computer program developed by Gordon and McBride (CET, Chemical Equilibrium with Transport properties). The GASP (GAS Properties) or WASP
(Water And Steam Properties) WASP} programs are implemented to obtain coolant thermodynamic and transport properties. Since the heat transfer coefficients of the hot gas and coolant sides are related to surface temperatures, an iterative procedure is used to evaluate heat transfer coefficients and adiabatic wall temperatures.


Figure 3: A rocket thrust chamber and nozzle subdivided into a number of stations

The temperature distribution within the wall is determined via an axial marching technique starting from station 1 to the last station. The program marches axially from one station to another. At each station a two-dimensional finite element model is used to determine the temperature distribution along the radial and circumferential directions. The axial heat conduction acts as internal heat source in the two-dimensional heat conduction model. When the axial march is completed, comparison is made between the results of the present march and that of the previous one to see if the convergence criteria in the axial direction has been met. If it is not met, the code starts again at the first station and makes another axial march. The process continues until convergence is achieved. A detailed description of this numerical model is outlined in the manual of RTE.

The following figures show some sample results of the rte (wall temperature distribution at various locations in the engine). Note that the temperature distribution is given for one cell and the indentation at the left is the cooling channel.

 Similar temperature distributions can be generated for all stations along the engine. In addition to the wall temperature distribution the program provides all transport and thermodynamics properties for coolant and combustion gases.

More detailed information on this program can be obtained from the following publications:
 

• Naraghi, M.H.N., ``RTE - A Computer Code for Three-Dimensional Rocket Thermal Evaluation," User Manual, 2002.
• Naraghi, M.H.N., ``RTE - A Computer Code for Three-Dimensional Rocket Thermal Evaluation,'' Grant NAG 3-892 report, prepared for NASA Lewis Research Center, 1994.
• Naraghi, M.H.N., ``RTE - A Computer Code for Rocket Thermal Evaluation," presented at the 1994 Thermal and Fluid Analysis Workshop, Cleveland, Ohio, August 15-19, 1994.
• Hammad, K.J., and Naraghi, M.H.N., ``Radiative Heat Transfer in Rocket Thrust Chambers and Nozzles,'' AIAA paper 89-1720, presented at the 24th AIAA Thermophysics Conference in Buffalo, New York, June 12-14, 1989; also AIAA Journal of Thermophysics and Heat Transfer, Vol. 5, No. 3, pp. 327-334, 1991.
• Naraghi, M.H.N., and Armstrong, E.S., ``Three Dimensional Thermal Analysis of Rocket Thrust Chambers,'' AIAA paper 88-2648, presented at the AIAA Thermophysics, Plasmadynamics and Lasers Conference, San Antonio, Texas, June 27-29, 1988.
• Naraghi, M.H.N., and DeLise, J., `` Conjugate Conductive, Convective and Radiative Heat Transfer in Rocket Engines," ASME publication HTD-Vol. 307, pp. 65-79, Preceding of the 30th National Heat Transfer Conference, Portland, Oregon, August 6-8, 1995.
• Delise, J.C., and Naraghi, M.H.N., ``Comparative Studies of Convective Heat Tranfer Models for Rocket Engines,'' AIAA paper 95-2499, presented at the 31st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, San Diego, CA, July 10-12, 1995.
• Naraghi, M.H.N., Quentmeyer, R.J. and Mohr, D.H., "Effect of a Blocked Channel on the Wall Temperature of a Regeneratively Cooled Rocket Thrust Chamber," AIAA paper 2001-3406, presented at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Salt Lake City, UT,  July 8-11, 2001

 

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